Omnivorous solar thermal thruster, cooling systems, and thermal energy transfer in rockets

ABSTRACT

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

INCORPORATION BY REFERENCE TO ANY PRIORITY APPLICATIONS

This application is based upon and claims the benefit of priority fromU.S. Provisional Patent Application No. 63/016,186 filed on Apr. 27,2020 and from U.S. Provisional Patent Application No. 63/159,957 filedon Mar. 11, 2021 and from U.S. Provisional Patent Application No.63/160,578 filed on Mar. 12, 2021. Moreover, Any and all applicationsfor which a foreign or domestic priority claim is identified in theApplication Data Sheet as filed with the present application are herebyincorporated by reference under 37 CFR 1.57. The entire contents of eachof the above-listed items is hereby incorporated into this document byreference and made a part of this specification for all purposes, forall that each contains.

BACKGROUND Field

The present disclosure relates to versatile rocket propulsion systems.For example, disclosed solar thermal rocket propulsion systems canoperate over a range of conditions and/or use a range of differentpropellants.

Related Technology

Related technology includes solar thermal rockets, which use solarenergy to heat propellants in a rocket chamber.

SUMMARY

A rocket engine (e.g., a solar thermal engine) can be provided,comprising: a thruster (e.g., a solar thermal thruster). The thrustercan be configured to: receive energy (e.g., solar energy) and one ormore propellants, and heat the one or more propellants using the (e.g.,solar) energy to generate thrust, wherein the thruster is furtherconfigured to adapt for a plurality of different propellant types,either singly or in combination simultaneously, and wherein the thrusteris further configured to use one or more of the propellant types in bothliquid and gaseous states.

The thruster can comprise: one or more regenerative channels configuredto direct flow of and simultaneously change thermal energy of the one ormore propellants.

The engine can be further configured to adjust a length and/or crosssectional area of the regenerative channels.

Each of the regenerative channels can have a fixed total length and cancomprise a plurality of inlet ports arranged along that length, and theengine can further comprise a plurality of inlet valves associated withthe inlet ports and configured to selectively inject the one or morepropellants into the regenerative channels via selected inlet ports suchthat a deployed length of the regenerative channels is adjustable.

Each of the regenerative channels can have a fixed total length, includea plurality of inlet ports, and the engine can further comprise aplurality of inlet valves configured to block, permit, or change flow ofthe one or more propellants into the inlet ports, thereby adjusting aflow of the one or more propellants through the regenerative channels.

The solar thermal thruster can comprise a pressure chamber and anexhaust cone, the pressure chamber connected to and axially aligned withthe exhaust cone; and the regenerative channels can follow a helicalpath around or through a surface of at least one of the pressure chamberand the exhaust cone.

The engine can further comprise: one or more electric thermal elementsconfigured to adjust a temperature of the solar thermal thruster; one ormore heat exchangers; and one or more valves to admit the one or morepropellants into the heat exchangers and adjust flows of the one or morepropellants to accommodate changing thermodynamic characteristics duringdifferent phases of rocket operations including startup, shutdown, andcontinuous operation at variable thrust levels.

The solar thermal thruster can comprise: a solar absorbing structurecomprising an assembly of partially reflecting, partially transmitting,and partially absorbing surfaces, thereby converting solar energy intothermal energy within the surfaces, and a transparent pressure-resistantwindow configured to transmit the solar energy into the solar absorbingstructure and contain the gases within the chamber.

The solar absorbing structure can be further configured to contain theone or more propellants in spaces between the surfaces such that thethermal heat energy is transferred to the one or more propellants viaboth thermal radiation and fluid conduction from the surfaces.

The solar absorbing structure of surfaces can be formed of one or moreof the following materials: metallic alloys and ceramics.

The solar absorbing structure can comprise a plurality of reflecting andabsorbing passages having triangular, rectangular, hexagonal, octagonal,or circular cross-sections, the reflecting and absorbing passages can befurther configured to provide a volume over which the solar energy isabsorbed simultaneously with providing a surface area and shape of asurface over which the propellant flows to absorb heat from the solarabsorber structure.

The solar absorber can further comprise a multi-surface light trappingsolar absorber comprising a honeycomb lattice, a bundle of thin walledtubes, or a coil of thin sheets.

The solar thermal thruster can further comprise one or more peripheralcooling channels, and the thruster adapts for different propellant typesby using at least one valve to adjust a deployed length of at least onecooling channel, thereby adjusting a cooling effect in the thruster.

The engine can further comprise: a solar collector configured to collectand redirect the solar energy; a solar absorber configured to absorb atleast a portion of the solar energy; and an adjustable solar fluxmodulator disposed between the solar collector and the solar absorber,the modulator configured to adjust thermal power input to the solarthermal thruster.

The solar flux modulator can comprise one or more of a variable geometryaperture, a shuttered opening, one or more blinds, and an opening atleast partially covered with a material having variable opticaltransmissivity.

The engine can further comprise: a store of cleaning propellant, whereinthe solar thermal thruster can be further configured to periodically usethe cleaning propellant to clean an inside surface of the solar thermalthruster and remove deposits made through the operation of the solarthermal thruster when using the one or more propellants.

The solar thermal thruster can be further configured to operate withdifferent combinations of depositing propellants and cleaningpropellants.

The window in the thruster can be configured as a lens to furtherconcentrate the sunlight into the thruster so as to increase the peaktemperature inside the thruster and increase the performance of thethruster.

An adjustable jacket cooling system can be provided for a rocket engine.The system can comprise: a combustion chamber having a strong peripheralwall and configured to contain propellant; a converging/diverging rocketnozzle at one end of the combustion chamber and configured to expelpropellant to produce rocket thrust; at least one cooling channel inthermal contact with the peripheral wall; at least one intermediateopening in the cooling channel; at least one valve configured to changea deployed length of the at least one cooling channel, thereby adjustinga cooling effect for the peripheral wall.

The cooling channel and valve can be configured to use the intermediateopening to change the structural cooling capability of the rocket enginesuch that it can operate effectively using at least two different typesof propellant (e.g., types that have significantly differentthermodynamic properties).

The adjustable jacket cooling system can be adapted for use with a solarthermal rocket engine. The engine can comprise a solar thermal thrusterconfigured to receive solar energy and one or more propellants, and heatthe one or more propellants using the solar energy to generate thrust.The adjustable jacket cooling system can allow the thruster to use aplurality of different propellant types in one or both of their liquidand gaseous states.

The cooling system and engine can have at least one cooling channel thatcomprises a regenerative channel configured to direct flow of andsimultaneously change thermal energy of the one or more propellants.

The cooling system and engine can be configured to adjust a deployedlength of the one or more regenerative channels. This can allow thesystem to be tuned for use with propellants having different properties.

A solar concentrator can be configured to power the solar thermal rocketengine, the solar concentrator comprising: a primary reflector; and asecondary reflector, wherein the primary reflector is configured toconcentrate the solar energy towards the secondary reflector, andwherein the secondary reflector is configured to reflect the solarenergy into a less converging, slightly diverging, or parallel beam, theintensity of the reflected solar energy being greater than the solarenergy prior to being reflected by the secondary reflector.

Two reflectors can be arranged in a Cassegrain configuration. In someembodiments, a Cassegrain configuration includes a paraboloidal primarymirror and hyperboloidal secondary mirror, can achieve similar resultsto those of a telephoto lens, and can have light brought to a focusthrough a perforation in the center of the primary mirror.

The primary reflector can comprise an orifice, and wherein the primaryreflector and the secondary reflector can be oriented such that thereflected solar energy passes through the orifice in the primaryreflector during normal operation prior to entering the solar thermalrocket engine.

The orifice can be positioned and configured to mitigate damagingeffects of pointing errors by rejecting the reflected solar energy whenincorrectly pointed.

A rocket engine system can comprise: a first propellant containerconfigured to hold deposit-forming propellant; a second propellantcontainer configured to hold deposit-cleaning propellant; a passageconnected to both the first and second propellant containers; a manifoldin fluid communication with the passage and configured to select orcombine propellant from the containers; and a cleaning control systemconfigured to control the manifold, thereby using propellants to reducedeposits within the rocket engine system.

The rocket engine system can further comprise: a third containerconfigured to hold cooling fluid; a passage configured to direct coolingfluid through structural elements of the rocket engine; and a coolingcontrol system configured to control the flow of the cooling fluid,thereby at least periodically cooling to maintain structural integrityof the rocket engine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a system configured to deliver concentrated solarenergy and use it in a solar thermal rocket.

FIGS. 2A-2H show how propellant injection tubes 7 can be orientedtangentially to a generally annular thruster chamber, thereby enhancingswirl of propellant and assisting in uniform propellant distribution.

FIG. 3A illustrates a portion of a solar thermal rocket system.

FIG. 3B illustrates thermal gradients within a rocket system.

FIG. 4 illustrates thermal shielding for a portion of a rocket.

FIG. 5 illustrates a system level overview of a solar thermal thrusterduring operation.

FIGS. 6A and 6B illustrate throttling of a solar thermal thruster usingan optical aperture.

FIG. 6C illustrates a protective hinge and cover plate that caninterrupt the formation of an energy beam.

FIGS. 7A-7C illustrate a solar absorber.

FIGS. 8A and 8B illustrate a solar absorber and incoming light.

FIGS. 9A and 9B illustrate actively cooling the engine to avoid failure.

FIG. 10 illustrates channels for vaporizing liquid propellant whileregeneratively cooling the material of the thruster.

FIGS. 11A and 11B illustrate a regenerative cooling and vaporizingchannel (e.g., during a startup sequence).

FIG. 12 illustrates a variable geometry channel's ability to modulatepropellant temperature.

FIG. 13 illustrates a variable geometry channel's ability to change massflow rate.

FIG. 14 illustrates an embodiment of an off an outward facing solarcollector.

DETAILED DESCRIPTION Examples of Rocket Propulsion Systems

A rocket can include a pressurizing chamber configured to contain apropellant gas that is heated to a relatively high temperature,typically above 1,000 K. The rocket can further include a nozzle with aminimum area throat through which the hot propellant gas is configuredto escape at a relatively high pressure, typically greater than 1 atmand sometimes are much as a few thousand atm, thereby providing thrustto propel the rocket forward. Aspects of this disclosure relate toimprovements to one or more of these structures.

A rocket can include at least the following two features: a pressurizingchamber configured to contain a hot propellant gas, and a nozzle shapedto allow the gas to escape. The rocket is configured to pass the hot,pressurized propellant gas from the chamber through throat of thenozzle, which can be shaped to use the gas to produce thrust. Forexample, the nozzle may be a converging-diverging expansion nozzle whichis configured to extract a flow of gas from the chamber and acceleratethe flow to a relatively high velocity (typically at least 100 m/s andsometimes up to 10,000 m/s, thereby producing thrust.

When hot gas is produced by the combustion of fuel and an oxidizer, thepressurizing chamber is usually termed a combustion chamber. In thiscase, the propellant may be a mixture of gaseous combustion products.When propellant is heated by non-combustion means such as electricalheating, optical heating, or other directed energy heating, then thepressurizing chamber may be termed a propulsion chamber. Often thepropellant is introduced into the rocket engine as a liquid and thenvaporized in the engine either by the heat of the chemical combustion orother means. In some implementations, the propellant can be any workingfluid, including a gas or a liquid that is used in a rocket engine.

Aspects of this disclosure relate to systems and methods configured touse concentrated solar energy to heat a propellant to a relatively hightemperature and pressure. This creates a “Solar Thermal Rocket” or STR.Throughout this disclosure, all reference to STR or solar thermal rockettechnology also applies to laser driven rockets. In laser driven rocketsa laser is be directed toward a rocket to provide analogous energyinputs described commonly throughout this disclosure as originating fromthe Sun. However, the same principles, structures, and solutions applywhere optical energy originates with a laser. For example, a laser maybe directed at a rocket to heat propellant and provide other energybenefits.

Aspects of the present disclosure also relate to improvement of alltypes of liquid propelled rocket engines, including electricalresistojets and arcjets, solar, nuclear, and chemical rockets, as wellas hybrids thereof, and the like. For example, all rockets that storepropellant as liquid can benefit from adjustable cooling channelgeometry, as described throughout this disclosure.

In aspects of the present disclosure, one or a combination of variouspropellant fluids may be used, including but not limited to: water,carbon dioxide, ammonia, hydrazine, or hydrogen. Propellant(s) may bestored in and introduced into the rocket engine either as gases orliquids. When introduced as liquid, the propellant is generallyvaporized by the engine, either internally in the chamber or prior toinjection into the engine. When liquid propellants are vaporized priorto their injection into the chamber, the device that performs thisvaporization may be referred to as a vaporizer. In general, significantenergy or power is provided to the vaporizer to allow the vaporizer toperform its function. In general, the physical geometry and the quantityof energy/power delivered to the vaporizer can be tailored to thepropellant fluid being used as each different propellant may vaporize ina different way depending on the physical and chemical properties of thepropellant.

It is anticipated that useful propellant materials will be extracted byspace mining processes that are applied on asteroids, moons, and/orother planetary bodies. Thus, a solar thermal rocket that can operateusing multiple alternative propellant materials is desired to enable thesolar thermal rocket to use available materials as propellant gas forpropulsion. Systems and methods for obtaining and using such propellantmaterials are described in U.S. Publication No. 2018/0051914, which isincorporated by reference herein for all that it contains.

In general, it is desirable for the temperature and pressure of gasinside the chamber of a rocket engine to be as high as possible becausethe speed at which the gas is ejected from the nozzle increases withincreasing chamber pressure and temperature.

Heat can be “lost” from the system. For example, heat can radiate orflow away from the chamber without contributing to thrust. Theusefulness of a rocket system can be increased by efficient conversionof the input thermal energy into produced thrust that avoids orminimizes such losses. It is desirable to minimize thermal energy lossesby reducing thermal heat flow out of the propellant gas and through thewalls of the propulsion chamber and the walls of the expansion nozzle.The heat loss problem may be increased when rocket thrusters are reducedin size. For example, smaller propulsion chambers have more surface areaper total volume than similar larger chambers. Thus, decreasing chambersize can lead directly to an increase in heat radiation losses and areduction in thermal efficiency.

It is often desirable to cool a portion (e.g., the exterior surface) ofthe chamber of a rocket engine to reduce the thermal stress on thematerials of the engine and/or to minimize heat leakage from the engineand therefore losses from the engine to the environment. Thruster wallcooling can be accomplished by flowing relatively cool propellantthrough channels in the wall of the thruster. This can result in“regenerative” cooling of the chamber because the propellant can absorbheat in this process (and, for example subsequently enter a combustionchamber with greater initial energy). If heat absorption is greatenough, the regenerative cooling channels also serve the purpose ofvaporizing the propellant and hence the regenerative cooling componentscan also function as the vaporizer of the thruster.

The systems and methods described herein provide improvements to thethermal efficiency and operational flexibility of solar thermal rocketsand can be applied to chambers of various sizes. Aspects of thisdisclosure can also simplify or otherwise improve the flow paths forpropellant gasses, thereby reducing construction costs and improvingoperating reliability. Aspects of this disclosure also relate tochanging the physical geometry of the vaporizer and the quantity ofenergy/power delivered to the vaporizer to improve the performance ofthe vaporizer and the thruster and allow the thruster to operate onalternative propellants.

The present disclosure may be applied to solar thermal rocket systems ofvarious sizes, dimensions, and proportions, using various propellantfluids (e.g., water, hydrazine, ammonia, argon, methane, hydrogen and/orpropane). Other materials (and/or the process that can produce them) aredescribed for example in U.S. Publication No. 2018/0051914 at FIGS. 22,23, 35, 38, 40, 41, 45, 50 and 51 (and their related descriptions).Features and advantages of the present disclosure will be apparent tothose of ordinary skill in the art from the above and from the followingadditional description, taken together with the accompanying drawings,in which like reference numerals refer to like parts.

An Omnivore™ thruster can use various materials as propellants. There ishigh leverage to having a propulsion system that can use propellantsthat require minimal processing in space prior to use. Some embodimentscan perform minimal processing to separate water from carbon dioxide andpurify the water for use as a propellant. Other embodiments can avoidprocessing and simply use all or a combination of fluid products of anoptical mining process (e.g., the processes disclosed in U.S.Publication No. 2018/0051914). The use of unprocessed fluids can beefficient and useful for obtaining propellants from asteroids, forexample, because it reduces the mass and complexity of any structuresand processes needed for distillation or separation of mined materials.The use of unprocessed or less processed fluids can also enhance theeffective productive yield of the propellant mining process itself. Forexample, in some embodiments the only processing may be filtering (notdistillation, for example). Filtering can help reduce the risk of largeparticles blocking a fluid nozzle, for example. Thus, raw, lessprocessed (e.g., filtered only), volatile products from an opticalmining process can be used to propel the solar thermal rocket.

U.S. Publication No. 2018/0051914 shows (see, e.g., FIGS. 6, 25-26, 49,53, and 57) various configurations for STRs; these figures and allrelated descriptions are among the materials incorporated by referenceherein.

In some example embodiments, a thruster can be formed from at least fourprimary parts: a thruster body (e.g., made of ceramic, metallic, orcomposite material), a transparent window (e.g., made of hightemperature glass, fused silica, sapphire, or other transparentmaterial) one or more clamps (e.g., Marman clamps), bonding connection,and seals to hold the window in place, and one or more ceramic foamsolar absorber inserts configured to absorb the solar energy andtransfer the absorbed heat to the working fluid. The thruster body maybe formed using 3D printing in some implementations.

A thruster body insert can be formed from ceramic, metallic, orcomposite material. The thruster body insert can be configured tofunction as a solar absorber and/or heat exchanger. In some embodiments,the solar absorber may be formed from a ceramic foam. The ceramicmaterials can be selected to be compatible with both SO2 and hydrocarboncontaminants because SO2 and hydrocarbon contaminants have been found inthe effluent of optical mining experiments. In certain embodiments, oneof the ceramic materials used for the thruster body can be Zirconia.Zirconia ceramics can be formed into a low density foam. Zirconia hasuses in the steel industry as a filter for cleaning debris from hightemperature molten metals. In some embodiments, another ceramic materialthat can be used for the thruster body is thorium oxide. Thorium oxidemay be a particularly useful material for flight applications. Thoriumoxide is a relatively high temperature oxidation resistant ceramic(melting point approximately 3300 degrees Celsius) that occurs innature. Thorium oxide can deliver specific impulse performance over 350s with water propellant, for example, and over 800 s with hydrogenpropellant. Thorium oxide is slightly radioactive and therefore bettersuited for use in space than in earth laboratory experiments performedby humans.

Design studies of three omnivore thrusters are further described inAppendix B to Provisional Patent Application No. 63/016,186, to whichthe present application claims priority (hereinafter “APIS”). Thesestudies explore the likely performance of omnivore thrusters usingmixtures of CO2 and H2O. Experiments showed that CO2 rich mixtures tendto produce higher thrust and efficiency, but lower specific impulse.

U.S. Pat. No. 4,036,012 provides additional background on STRpropulsion, thruster performance potential, inflatable solarconcentrators, liquid hydrogen thermal control, and propellant feedsystems. U.S. Pat. No. 4,036,012 is also incorporated by referenceherein for all that it contains.

FIGS. 1-4 of U.S. Pat. No. 4,036,012 show STR embodiments andimprovements that can be used with the structures and methods describedin U.S. Publication No. 2018/0051914, APIS, and U.S. Pat. No. 4,036,012.These non-limiting examples illustrate principles and structures thatcan apply to all other STR embodiments disclosed herein (and in U.S.Publication No. 2018/0051914, APIS, and U.S. Pat. No. 4,036,012).

FIG. 1 illustrates a system configured to deliver concentrated solarenergy to a solar thermal rocket. The system includes one or morereflecting surfaces 2, a transparent window 4, a pressurized propulsionchamber 5, a propellant supply system 6, and an injector channel 7.

Referring to FIG. 1, solar energy 1 from the sun is reflected andconcentrated by the one or more reflecting surfaces 2. In someembodiments, the one or more reflecting surfaces 2 may be curved asillustrated. The converging solar energy 3 passes through thetransparent window 4 into the pressurized propulsion chamber 5. Apressurized propellant gas can be simultaneously delivered from thepropellant supply system 6 through the injector channel 7 into thepressurized propulsion chamber 5.

As shown in FIG. 1, the injector channel 7 can be positioned inproximity to the window 4 to provide a degree of cooling for the windowmaterial. In some embodiments, the window 4 can be fabricated from ahighly temperature tolerant material such as fused quartz or sapphire.

An STR can position a thruster at or near the focus of a solarconcentrator when in use. Advanced 3D printing of high temperatureceramics (e.g., Zirconia foam) can be used to form a monolithic thrusterbody (e.g., in the right shape to occupy or fill a thruster chamber andhaving a low density closed-cell ceramic foam structure). A window(e.g., sapphire) can provide optical access to an enclosed thrusterchamber. One or more propellant inlets can be located at or near thewindow to cool the window. The ceramic foam structure can be a solarabsorber and provide a heat interchange as propellant flows through opencells of the foam structure and solar energy heats the cell walls,thereby allowing the walls to transfer heat energy to adjacentpropellant in the cells. This increases the pressure in the chamber andcauses flow toward the nozzle's throat. One or more (e.g., additional)propellant inlets can enter the sidewalls of the thruster body near thenozzle's throat and flow toward the top of the foam structure usingannular channels that open near the window. In a process of regenerativeheat capture, this fluid can absorb heat radiating from the chamber intoits walls and therefore arrive in a warmer state than fluid that has notpassed through these channels in the thruster body. By capturing theheat in this manner, the regenerative cooling channels increase theefficiency of the thruster by returning heat, which would otherwise belost, to the propellant prior to its entry into the chamber.Accordingly, the channels can also be considered regenerative heatingchannels in the sense that they also productively pre-heat propellant.The thermal energy transfer they facilitate is thus beneficial in manycomplementary ways. Cooling chamber walls prolongs their life;pre-heating propellant increases thrust or makes the chamber moreefficient; providing for customized or dynamic channel lengths can allowthe thruster to be tuned for use with a wider variety of propellants;and the overall system wastes less energy, achieving a greater netefficiency over the longer period of time it survives. Thecooling/heating channels have a synergistic effect and can make anyrocket engine more durable, efficient, powerful, versatile, andproductive.

In some embodiments, the window can be made of a flat transparentmaterial. In aspects of the present disclosure, the window can be madeof lens shaped transparent material convex on one or both sides suchthat the light is concentrated into the center of the thruster toincrease peak temperature and therefore performance of the thruster.

Most common gas molecules are substantially transparent to the dominantenergy wavelengths contained in sunlight. It follows that theconcentrated solar energy 3 couples only weakly to most gasses. Aspectsof this disclosure can overcome this potential limitation in the solarthermal process by introducing an intermediate thermal conversion step.

In certain implementations, the propulsion chamber 5 can contain aquantity of open-cell porous ceramic material 8. The ceramic material 8may be partially transparent and partially absorbing to solar energy inits bulk structure. The ceramic material 8 can also contain a pluralityof open channels through which solar energy can penetrate deeply intothe structure of the ceramic material 8. The open-cell structure of theceramic material 8 may further permit the flow of propellant gas throughthe ceramic material 8. The use of such a ceramic material 8 can enableconcentrated solar energy to be absorbed efficiently in heating theceramic material 8. Simultaneously, the large internal surface area ofthe porous ceramic material 8 can efficiently transfer heat to thepropellant gas by direct conduction.

The pressurized propellant gas passes from the injector 7 through theceramic material 8 where the propellant gas is heated to a relativelyhigh temperature and then into the converging-diverging expansion rocketnozzle 9. As in most rockets, the nozzle 9 converts the random thermalmotion of the high temperature gas into a directed high velocity flow.The nozzle 9 can extract a flow of gas from the chamber and acceleratethe flow to a relatively high velocity, thereby producing efficientthrust.

It is desirable to heat the propellant gas to the maximum possibletemperature achievable by the rocket. In certain embodiments,non-imaging anidolic optical systems can be used to heat the ceramicmaterials 8 and infuse the propellant gas within the ceramic totemperatures above 2000 K (degrees Kelvin). Such temperatures can beachieved with concentrating systems 2 having a collection area equal toor greater than 1 square meter when located at a distance from the sunsubstantially equivalent to the orbit of earth. Embodiments of the solarthermal rockets disclosed herein can be used throughout the inner solarsystem and to beyond the orbit of Mars.

FIGS. 2A-2H show how propellant injection tubes 7 can be orientedtangentially to a generally annular thruster chamber, thereby enhancingswirl of propellant and assisting in uniform propellant distribution.Holes 14 located around the circumference of the chamber can communicatewith propellant inlet tubes 7 and allow primary and regenerativelyheated propellant to flow into the chamber. A window 4 made of sapphireor other high temperature capable transparent material can be attachedwith high temperature O-rings, gaskets, c-rings, and/or Marmon clamps,for example. In the thruster chamber, structure (e.g., a step ledge) 15can be provided to support a solar absorber insert (e.g., one formedfrom ceramic foam) 121. FIGS. 2F and 2H show a cross-section view of thestep ledge 15 and an isometric cutaway of a long chamber design allowingdifferent foam inserts 121 for experimentation. Advantageously, anomnivore thruster can use “contaminated” (or non-pure) propellantshaving various constituents and use the direct products of opticalmining. Although large debris particles (e.g., more than 0.4 mm indiameter—depending on foam cell size), may clog the foam solar couplerand may therefore be filtered out, it is advantageous to not requireadditional processing, in various embodiments.

FIG. 3A shows a cross-section view of the solar thermal propulsionchamber 5. The transparent window 4 is captured in a top plate 11 whichprovides a pressure seal against the internal gas pressure. Thepropulsion chamber wall 5 is shown with a substantial thickness in crosshatched representation. The wall 5 can be fabricated from a hightemperature ceramic material such as Zirconia, for example. The entirerocket nozzle can be contained in a metal jacket 12. In someembodiments, an intervening distance between the jacket 12 and the wall5 can be open to ambient space. The gas injector channel 7 is shown topenetrate both the metal jacket 12 and the propulsion chamber wall 5.

FIG. 3B shows in simplified detail the temperature gradients in andaround the solar thermal thruster. Within the propulsion chamber 5, itcan be desirable to keep the gas temperature high. In FIG. 3B, thetemperature is shown as about 2000 K, but this is just an example. Theouter, generally cylindrical, metal jacket 12 and the top plate 11 canbe at nearly room-temperature of approximately 300 K. Such a simpledesign can present a severe thermal gradient across the Zirconia ceramicin the chamber wall 5. Thus, with this configuration, there may be arelatively high thermal stress across the chamber wall 5, which canresult in a high probability of fracture failure.

FIG. 4 shows a design element of aspects of this disclosure that canrelieve the potentially damaging thermal stress described in connectionwith FIG. 3B. An additional shielding feature can be added (for example,a thin cylindrical metal foil 13 as shown in FIG. 4) between theZirconia ceramic of the propulsion chamber wall 5 and the outercylindrical metal jacket 12. The foil 13 can serve as a radiationshield. During steady state operation, the foil 13 may come to anintermediate temperature between the propulsion chamber wall 5 and themetal jacket 12. By using the foil 13 as shown in FIG. 4, heat radiationaway from the propulsion chamber wall 5 can be substantially reduced.Thermal stresses can be reduced and the internal temperature of thepropulsion chamber 5 can be maximized, thereby increasing rocket thrust.

One or more cylindrical foils 13 (or other thermal or radiationbarriers) may be added. In some embodiments, they may be similarlyspaced apart. When operating in space, the nozzle 9 propels high speedgasses directly into a vacuum. Similarly, the space between thepropulsion chamber wall 5 and the outer cylindrical metal jacket 12 canalso be open and therefore evacuated to the very low pressure of deepspace. Interposing one or more barriers (e.g., foil walls 13) enhancesthe insulation properties of the vacuum.

Some solar thermal rocket embodiments attempt to reduce or moderatetemperature gradients in a propulsion chamber wall 5 by circulating coolpropellant gas through channels around the propulsion chamber wall 5.See FIGS. 2A-2E. Aspects of the present disclosure simplify and reducecost of manufacturing when compared to implementations that circulatecooling propellant gas around the propulsion chamber wall 5. Simplerinlet passages for propellant(s) such as those described herein can alsoreduce the risk of being clogged by debris.

Further Examples of Rocket Propulsion Systems

Rocket engines can operate on a single propellant or a mixture ofpropellants. Aspects of this disclosure relate to solar thermalpropulsion rocket engines designed to operate in space at variable powerand thrust levels and capable of operating with various propellants usedsingly or in combination simultaneously. The propellants can be storedas either liquid or gaseous fluids.

For spacecraft propulsive applications in Earth orbit, cislunar space,or other orbits, it is advantageous to use readily available solarenergy, or sunlight, as the primary power input to the propulsionsystem. This permits the extraction of propulsive energy from theambient environment rather than carrying that energy in the form ofchemical propellants or via extraction of propulsive energy from anelectric power source which could be energized by solar arrays, nuclearenergy, or batteries. One advantage of direct solar thermal propulsionis that the energy source need only be a lightweight reflector thatconcentrates the solar energy onto the thruster or engine. Applicationsof solar thermal propulsion systems include but are not limited tospacecraft attitude control, station keeping, orbit raising andlowering, orbital plane changes, and more generally transfers to andfrom asteroids, lunar orbit, and orbits around planets. Solar thermalpropulsion systems can be used wherever there is sufficient sunlight, asmay be found throughout the inner solar system out to approximately theorbit of Jupiter.

In situ resource utilization (ISRU), is the process of using materialsharvested in space to supply space operations. ISRU can be used toenable more pervasive and cost-effective space travel. With ISRUspacecraft propellant may be harvested, manufactured, and stored toprovide in-space refueling. Mid-mission refueling reduces launch massand therefore cost. U.S. Patent Publication No. 2018/0051914 providesexample systems and methods for ISRU. ISRU can reduce or eliminate theneed to launch propellant from the surface of the Earth to supply spaceoperations. ISRU may therefore allow increased payload transport for thesame launch cost.

The solar thermal propulsion systems that utilize propellants collectedvia ISRU described herein represent an improved and less costlyalternative to chemical propulsion systems for near Earth missions andoperations within the inner solar system. To fully capitalize on thispotential, it is advantageous for such rocket engines to be flexible and“omnivorous” enough to consume whatever propellants are available fromin-situ sources, with the same engine being able to operate on differentpropellants at different times. Aspects of this disclosure are motivatedby the desired for propellant flexibility. Likewise, as the spacecraftmoves to different locations in space, the intensity of sunlight mayvary, and so it may be further advantageous for the solar thermal rocketto be able to change its operating power in accordance with theintensity of sunlight available. Finally, there may be cases in which itis desirable for the thrust level or specific impulse of the rocket tochange. Aspects of present disclosure are designed to satisfy one ormore of the above-indicated problems for operational flexibility.

The components of a solar thermal propulsion system may include apropellant vaporizer, a solar concentrator, a restricting aperture, anda solar absorber in addition to other elements commonly found in rocketsystems. The use of multiple propellant types, variable thrust level,and the ability to operate at variable solar input power levels caninvolve added functionality as discussed in at least the five examplecases below.

First, as is also the case with more typical chemical propulsion systemsor rocket engines, the propellant gases in the subsonic pressure chambermay be operated at a greater temperature than the practical workingtemperatures of the material structure of the rocket engine. Hence, itis desirable to manage the heat load of hot gases in a solar thermalrocket to avoid structural damage to the engine.

Second, it is often desirable to inject the propellant fluid into thepressure chamber in the gaseous phase. To do this, the intrinsicheat-of-vaporization can be supplied prior to injection into thepressure chamber. Furthermore, when liquid phases of propellants areconverted to vapor, they can be accompanied by a substantial increase inpressure for a given closed volume. If sudden unsteady phase changesoccur in the pressure chamber of a solar thermal rocket engine, theresultant stresses and temperature can spike, particularly duringtransient operations such as startup and shut down, which may causefailure of structural components or reduce the efficiency of the engine.It is desirable to minimize the deleterious effects of such unsteadyoperation due to uncontrolled phase changes. Hence, in some embodiments,heat can be applied to first vaporize the liquid into gas prior toinjection to at least partially address the unsteady operation. Then,further heat can be applied to increase the gas temperature in therocket pressure chamber. Increased chamber temperatures may be directlyrelated to increased propulsive specific impulse, which can be a keyfigure of merit for many space propulsion systems.

Third, it is desirable for rocket engines which are designed to usemultiple propellant types in flight to be able to dynamically vary theheat flux to the propellant vaporization process and the mass flow rateinto the pressure chamber in order to adapt the propulsive andthermodynamic characteristics in accordance with each propellant type.

Fourth, when focusing solar energy onto the absorber in solar thermalrocket engines, it is desirable to provide additional mechanisms todynamically adjust the solar heat flux to the engine. In someembodiments, heat flux may be adjusted during transient operations suchas startup and shutdown or for the purpose of controlling system thrustand specific impulse during normal operation.

Fifth, effective absorption of solar radiation and heat transfer to thepropellant may be important considerations for the overall design of anefficient solar thermal propulsion system. Efficient use of heat canreduce the size of the solar collector needed to collect adequate solarenergy and can reduce the size of any required waste heat radiators andcan be used to obtain the maximum possible operating fluid temperaturesin the engine. Maximum operating temperature is desirable as itincreases the maximum possible specific impulse and thrust of theengine.

It is desirable for the omnivore thrusters described herein to be ableto work using a wide variety of propellant combinations. Somepropellants, such as methane or other hydrocarbons, if used in theomnivore thruster, may deposit materials inside the thruster which overtime could degrade the thruster's performance. When hydrocarbons formsuch deposits in a thruster or rocket engine, the process is generallyreferred to as coking, but deposits can be caused by a variety ofdifferent chemical processes and can involve a variety of differentmaterials. Because a given omnivore thruster can operate on differentpropellants at different times, the thruster can mitigate the problem ofcoking or other deposition processes by alternatively using differentpropellants. These propellants can be selected to remove depositsremaining from previously-used propellants, effectively cleaning theinside of the thruster while simultaneously propelling the thruster. Forexample, running the thruster using water can have the effect ofcleaning any carbon material that might be coked onto the inner surfacesof the thruster due to operation with hydrocarbon propellants. Thus, acontrol system can be used to select a series of propellants where atleast one subsequent propellant can cure or mitigate (e.g., clean)harmful effects (e.g., deposits) from previously used propellants.Alternatively, a combination of propellants can be used tosimultaneously clean or prevent deposits. For example, a mixture ofdeposit-forming propellant and deposit-cleaning propellant can be usedto alleviate problems from deposits.

Further Examples of Solar Thermal Rockets Configured to Use MultiplePropellants

FIG. 5 illustrates a solar thermal rocket using mirrored collectors 118to redirect and concentrate sunlight 120 into a high intensity beam 103.In particular, the solar thermal rocket of FIG. 5 includes the mirroredcollectors 118, an optical aperture 102, and a solar thermal thruster106. The mirrored collectors 118 are configured to direct the light 103at the solar thermal thruster 106 through the optical aperture 102. Thesolar thermal thruster 106 includes a solar absorbing medium 119configured to absorb energy from the light received via the opticalaperture 102.

Typical rocket engines employ high energy density within their pressurechambers to create desired flow characteristics and thrust. For example,chemical rocket engines use the potential chemical energy stored in themolecular bonds of rocket fuel to concentrate energy within thecombustion chamber.

In the solar thermal rocket of FIG. 5, sun light 120 is collected andconcentrated by the one or more reflective surfaces 118 to form the highintensity beam 103. The beam 103 is absorbed in the absorber medium 119located within the thruster body 106. The absorber 119 is configured tobe heated by the beam 103 to a high temperature at which the absorber119 can transfer heat energy to a surrounding gaseous propellant. Byheating the absorber 119 to a sufficiently high temperature, theefficiency of the conversion of the propellant into thrust can beimproved.

FIGS. 6A and 6B illustrate the use of an optical aperture 102 (which canbe adjustable, for example) as an energy control mechanism. Inparticular, FIG. 6A illustrates the adjustable optical aperture 102 in alarger configuration 102A configured to allow light to pass into thethruster body unobstructed, and FIG. 6B illustrates a smallerconfiguration 102B for an adjustable optical aperture 102. Compared to102A, this smaller configuration reduces the amount of light allowed topass.

The ability to modulate a thruster's 106 input energy is desirable inspacecraft operations. Startup operations use a slowly increasing energyinput to prevent damaging pressure or temperature transients, which maydamage components of the thruster. When running at nominal power level,it may be desirable to change the type of propellant without shuttingdown and restarting. In general, each type of propellant may have acharacteristic operating temperature and heat of vaporization. Thus, itis desirable to compensate for changes in the energy input to thethruster 106. An adjustable optical aperture 102 can open fully 102A asshown in FIG. 6A for maximum unobstructed power flow 104 or can beconstricted 102B to reduce incoming power 105. In some embodiments, thediameter of the adjustable optical aperture 102 can be adjusted to oneof a plurality of predefined values, or can be adjusted in a steplessfashion to any desired diameter. As an alternative to, or in additionto, an adjustable optical aperture, a shutter system can be used tocontrol or adjust light or heat. For example, U.S. ProvisionalApplication No. 63/055,231 shows an example hinge and shutter system(e.g., FIG. 6) which is also illustrated in FIG. 6C herein.

FIG. 6C illustrates a protective hinge and cover plate that caninterrupt the formation of an energy beam. Referring to FIG. 6C, a hinge224 and cover plate 225 are affixed to cover the front surface ofreflecting element 213. In this configuration, converging rays fromcurved reflector 212 are prevented from reflecting and combining to formenergy beam 216. Depending on the surface reflectance of cover plate225, the incident light will be partially absorbed and partiallydiffusely reflected. The camera 219 will image a diffuse glow upon thesurface of cover plate 223. The image position of the glow spot on thesurface of the cover plate indicates the accuracy of alignment of curvedsurface 212 to the incoming solar energy 215 without requiring areflection from element 213. In this manner, curved surface 212 may beaccurately aligned to the direction of the sun before a powerful energybeam is directed further into the spacecraft. Furthermore, the curvedsurface may be accurately aligned in the direction of the sun withoutneed for a separate sun tracker or sun observing device.

Referring again to FIG. 6C with the cover plate 225 in the blockingposition as shown, a significant amount of solar energy 215 can beabsorbed by the cover plate 225 thereby causing its temperature to rise.As the plate 225 heats, it will radiate long wavelength thermal energyin all directions. Nearby structures, such as lenticular structure 210with transparent upper surface 211 will intercept a portion of theradiated thermal energy and also begin to warm. In this manner,lenticular structure 210 and other surfaces can be warmed as needed toprevent moisture or ice build that could potentially degrade its opticalperformance. Since moisture may be produced during asteroid miningoperations, it is desirable to heat optical surfaces as needed.

FIGS. 7A-7C illustrate different embodiments of a solar absorber whichcan be used in the solar thermal thruster 106. For example, FIG. 7Aillustrates a cross sectional view of a solar thermal thruster 106 witha solar absorber 119. FIG. 7B illustrates a cross section of ceramicfoam solar absorber 121. FIG. 7C illustrates a multi-surface lighttrapping solar absorber 113, which may be embodied as a honeycomblattice, a bundle of thin-walled tubes, or a coil of thin sheets.

This disclosure hereby incorporates by reference U.S. Pat. No.5,138,832A to Pande, which describes a method of absorbing concentratedsolar energy in a ceramic open-cell foam-like materials consisting ofmany small hollow cells.

However, ceramic foam absorbers have several disadvantages compared tosome systems disclosed herein. First, the concentrated solar energytends to be deposited substantially in the surface layers of the foamrather than throughout the bulk volume leading to localized excessiveheating. Second, the small cell sizes of ceramic foam absorbers 121represent considerable mechanical impedance to the flow of heat-exchangefluids (e.g., propellant) and can produce large undesirable pressuredrops. Third, the small hollow cell walls of ceramic foam absorbers 121tend to be fragile when subjected to thermal and mechanical stressresulting in small particle shedding and uncontrolled debriscontamination to the downstream rocket nozzle. Fourth, the foamstructures of ceramic foam absorbers 121 are fragile and subject tofracture and breakup due to thermal and mechanical stress duringoperations.

In some embodiments such as that of FIG. 7C, the solar absorber includesa multi-surface light trapping solar absorber 113. In some embodiments,the multi-surface light trapping solar absorber 113 can include ahoneycomb structure array of tubes 113 (also referred to as channels).The structure 113 can be fabricated with dimensions configured to allowfor more accurate control of energy absorption than the semi-randomstructure of a foam, such as a ceramic foam absorber 121. The incomingsunlight 103 can be reflected multiple times by the walls of the tubes113 of the honeycomb structure, where the sunlight 103 is substantiallyabsorbed. Depending on the implementation, the individual tubes of thestructure 113 may have different cross-sectional shapes, for example,the tubes 113 may have triangular, rectangular, hexagonal, octagonal, orcircular cross-sections.

Devices configured to substantially absorb large amounts of opticalpower are generally referred to as light traps or beam dumps. Theoperational principle of these devices is to reflect a beam of incominglight from multiple surfaces at nearly grazing or similar low angles ofincidence. Reflections are designed to be specular rather than diffuseto minimize light backscatter in the direction of the incoming lightsource. Light may be partially or strongly absorbed and partiallyreflected at each surface. After many reflections, the incoming light issubstantially absorbed in total. The multi-surface light trapping solarabsorber 113 can be configured to act as a light trap or beam dump inorder to absorb substantially all of the received sunlight 103. Oneadvantage to the use of a multi-surface light trapping solar absorber113 is that heat energy is absorbed over a relatively large absorptionarea or volume compared to other implementations.

This disclosure hereby incorporates by reference U.S. Pat. No. 5,214,921to Cooley, which employs a multiple reflection solar energy absorptiontechnique to concentrate solar energy for the purpose of heating bothliquid and gaseous fluids. The physical principles of Cooley can beapplied to heating propulsion fluids for rocket applications.

For the solar thermal rocket application, it is desirable that theabsorbing materials maintain their structural strength at the relativelyhigh temperatures used to achieve peak thrust of the rocket. It isuncommon to find materials suited for this extreme temperatureapplication that also have the preferred high optical absorbingproperties. By using light trap embodiments (e.g., by using themulti-surface light trapping solar absorber 113) in the presentdisclosure, the reflecting surfaces are not required be as stronglyabsorbing because of the many reflection opportunities. For example, ina light trap configuration (e.g., the multi-surface light trapping solarabsorber 113), a material that is a specular reflector with areflectivity of about 80% corresponding to an absorptivity of about 20%can produce an effective absorptivity for the solar absorber of morethan 90%. Thus, aspects of this disclosure are able to use a widervariety of materials for constructing the multi-surface light trappingsolar absorber 113, which is a distinct advantage over other systems.

FIG. 8A illustrates a cutaway of an operating thruster 106 with ahoneycomb of solar absorber tubes 113. An absorber can be located in allor a portion of a thruster 106. FIG. 8B illustrates the path 114 oflight 103 in a single tube 113 as light is absorbed. In particular, aportion of the light 103 may be absorbed and a portion of the light 103may be reflected as the light 103 travels through the tube 113 along thepath 114.

FIG. 9A and 9B illustrate that cooling the body of a thruster isdesirable for efficient long-life operation of a solar thermal rocket.In particular, FIG. 9A illustrates a thruster 118 that has failed due tolack of sufficient cooling, and FIG. 9B illustrates a thruster 106operating normally with cooling.

Higher temperatures in the pressure chamber of a rocket generally relateto higher rocket specific impulse. One limiting factor on pressurechamber temperatures is the thermal capabilities of available materials.By cooling the walls of the thruster 106 as illustrated in FIG. 9B(e.g., by one of various disclosed mechanisms, including absorption,dispersion, heat transfer, fluid circulation, material diversity,regenerative cooling, etc.), heat can be continuously removed,dispersed, or re-allocated, allowing for higher internal propulsive gastemperature without material failure. Without a cooling mechanism,attempting to operate above safe temperatures may result in overheatingand failure 118 as shown in FIG. 9A.

FIG. 10 demonstrates vaporizing liquid propellant using regenerativecooling. In particular, FIG. 10 illustrates a design of the thruster 106which enables a liquid propellant to flow through channels 109 (alsoreferred to as vaporizer channels, heating/cooling channels orregenerative channels), where the liquid propellant can absorb energy108 from the gases in the pressure chamber 107. This energy 108 cancause a phase change of the propellant from liquid to gaseous beforeinjection into the pressure chamber 107. The pressure chamber 107 isalso connected to an exhaust cone 123. The exhaust cone 123 may beconnected to the pressure chamber 107 along an axial direction of thethruster 106.

It is desirable to vaporize liquid propellant before injection into thepressure chamber 107 to minimize damaging pressure spikes within thechamber which can be caused when liquid droplets suddenly vaporize athigh temperature. In the embodiment of FIG. 10, propellant is configuredto flow through the cooling channels 109, where the propellant absorbsheat 108 though the walls of the chamber which are heated by the hotgases in the pressure chamber 107 thereby causing a phase change fromliquid to gas in the propellant. After traveling through the coolingchannels 109, the propellant gas is injected into the pressure chamber107. The configuration of the pressure chamber 107 can be referred to asan integrated vaporizer.

Channels 109 can be configured to pass through the walls of a thruster106, for example, the walls of a pressure chamber 107. In someembodiments, such channels can be joined to form a continuous channel,allowing fluid to pass progressively through a series of coils or turnsaround a chamber 107. The channels can be configured to increase surfacearea and/or time fluid is in contact with heated side walls, to maximizeenergy transfer opportunity. The structure of the channels can beconfigured to engineer flow characteristics—for example speed. Channeldesign can allow for continuous flow despite potential blockingmaterials in some portions of a channel. Channels can be formedintegrally within side walls to enhance fluid flow (both within thechannels and across smooth sidewalls outside the channels. Channels canpass laterally and/or longitudinally, in varying combinations.

Solar thermal heating has several advantages over electrically heatedthermal rockets. Solar thermal rockets may use solar radiation todirectly heat the heat-transfer surfaces. Electric heating can involveseveral conversion steps from solar cells to electric power managementsystems to resistive heating of heat transfer surfaces and is,therefore, relatively less efficient than direct solar thermal.

Although not fully illustrated in FIG. 10, the regenerative channels 109may have a helical form wound around the surface of the chamber or maybe manufactured into the pressure chamber 107 and the exhaust cone 123.

FIGS. 11A and 11B illustrate embodiments of a time sequencerepresentation of the channels 109 during a startup sequence. FIGS. 11Aand 11B respectively depict two embodiments of a startup sequence havingtwo stages. In first stage, the thruster is brought to operatingtemperature. The second phase minimizes transient effects ofliquid-to-gas phase change.

In more details, FIG. 11A illustrates a propellant flowing through thechannels 109 during a startup sequence which reduces unwanted fluiddynamic transients. The large arrows indicate the passage of time duringthe transition for startup to normal operations. FIG. 11B illustrates anembodiment of a startup sequence, electric heating elements 122 raisethe temperature of the thruster 106 to operating before propellant flowand concentrated sunlight input.

In a system in which fine pressure control is desired, unpredictabletransient fluid dynamics are undesirable. In the embodiment of FIG. 11A,the transient effects of liquid-to-gas phase change can be reduced orminimized by beginning with a gaseous propellant 110 at startup. Liquidpropellant may be slowly added while reducing gaseous input 111 untilthe input is fully liquid 112. The rate of change of the inputgas-to-liquid ratio can be controlled to keep the system ininstantaneous steady state throughout the startup transition to avoidexcessive heating of any component of the system during transientoperation.

As the thruster 106 can be designed for intermittent use, the thruster106 will cool during periods of inactivity. This thermal cycling hasnegative repercussions for all rocket engines. In the case of a solarthermal thruster as disclosed herein, a cold start may present a uniqueproblem when using liquid propellant. In order for the regenerativecooling channels 109 to vaporize the liquid propellant as depicted inFIG. 10, the thruster body 106 may need to be at or near operatingtemperature. Heating the thruster 106 with concentrated light withoutflowing propellant can risks thermal damage to one or more components ofthe thruster 106 because the flowing fluid can be used to maintainthermal equilibrium of the thruster to avoid over heating components.

FIG. 11A shows how the thruster 106 can be brought to operatingtemperature using concentrated light while a gaseous propellant flowsthrough the vaporizer channels. Thermal damage can be mitigated in thisembodiment as the gaseous propellant removes heat, allowing the systemto reach thermal equilibrium under operating conditions.

FIG. 11B shows a thruster 106 that includes an electric heater 122configured to heat the thruster 106 to operating temperature in theabsence of propellant flow and concentrated light. A control loop suchas a proportional-integral-derivative (PID) loop or other control logicsystem and accompanying temperature sensors can be used to provide aspecific thruster 106 temperature to be reached and maintained duringstartup.

In the embodiment of FIG. 11B, liquid propellant flow can be graduallyincreased until reaching operating levels, which may be solelyelectrically heated. At this time in the startup sequence, the thruster106 can function as an electro-thermal thruster. Electro-thermalthrusters, generally referred to as resisto-jets, can use electricity toheat propellant and have been a flight proven technology since the1960s. After thermal and fluid dynamic steady state are reached, thethruster 106 can transition from an electro-thermal mode tosolar-thermal thruster mode by simultaneously reducing power to theheating elements 22 and increasing input concentrated light until thethruster 106 is operating solely on light.

As described in FIG. 10, electrically vaporizing liquid propellants foruse in a solar thermal rocket may be relatively inefficient. However,the length of the startup sequences described herein may be relativelyshort compared to the total runtime making any slight inefficienciesacceptable.

FIG. 12 illustrates fluid flow resulting in controlled propellanttemperatures. In particular, FIG. 12 depicts a method of controlling thetemperature of propellant heated by the regenerative cooling channels109 described herein. It is desirable in a complex system to correlateor match physical variables with controlled degrees of freedom. Variablefluid properties of the propellant which can be controlled, for example,include temperature and mass flow rate. Some fluids, notably water, havea relatively large heat-of-vaporization, while others, such as ammoniahave somewhat smaller heat-of-vaporization. For this reason, it wouldnormally not be possible for a regeneratively cooled thruster withchannels for vaporizing the propellant to operate using both water orammonia, for example. However, a thruster can be designed to work onboth propellants or differing mixtures of them at different times. Thiscan be accomplished, for example, using variable geometry channels. Theflow geometry may be varied through the use of valves 115 and 116.Channels can have a shape that varies over time and/or that varies alonga length or other dimension thereof.

FIG. 12 illustrates a plurality of cooling channels 109, an upstreamvalve 116, a downstream valve 115, and the thruster 106. Additionalvalves and channels can be used, serially or in parallel (see FIG. 13).The system is configured to modify and control the propellanttemperature by injecting (or allowing flow of) a variable proportion ofthe propellant downstream through the downstream valve 115 with respectto an amount of the propellant that is injected through the upstreamvalve 116 of the cooling channels 109. Propellant injected upstream(e.g., through valve 116) interacts for a greater time and distance withchannel walls, absorbing more energy in the process. Propellant injecteddownstream (e.g., via the downstream valve 115) has less interactionwith the hot channel walls than propellant similarly injected upstreamat valve 116, and thus receives less energy 108 from the walls of thecooling channels 109. Increasing flow through downstream valves (e.g.,the valve 115)—for example, by opening more of them, or opening them toa larger degree—while simultaneously reducing flow through upstreamvalves (e.g., the valve 116) can concentrate a thermal transfer (e.g.,cooling) effect in an area near the downstream channel 109. This canreduce the overall propellant temperature at the input to the thruster106, for example.

FIG. 13 illustrates fluid flow that allows for increased mass flow ratesat controlled flow velocity. In particular, FIG. 13 depicts a systemincluding a plurality of metering flow valves 117, a plurality ofcooling channels 109, and a thruster 106. The system is configured tocontrol the mass flow rate of propellant through the regenerativecooling channels 109. In this embodiment, multiple channels 109 areplaced in parallel. The metering flow valves 117 can be opened or closedin greater or lesser numbers or to greater or lesser degrees and arethereby configured to direct more or less propellant through the flowchannels 109 to control the mass flow rate. By using a greater number ofchannels 109, the system can achieve higher mass flow rates withoutincreasing fluid velocity. It is desirable to control fluid velocity,for example, to control the amount of time the propellant is present inthe flow channels 109, and thus, the amount of thermal interaction time(e.g., during which the propellant is heated via contact with walls ofthe flow channels 109). With excessive velocity the propellant has lesstime to interact with the hot channel walls causing an undesireddecrease thermal energy transfer to the propellant.

The channels described in FIG. 12 and FIG. 13 can have various shapes,lengths, configurations, and cross sections. They can be independent orintersect. They can have multiple valves and openings, they can beformed from various materials, and they can have consistent or varyingcross-sections (within, between, or among them). In some embodiments, athruster wall can have (e.g., contain or be in thermal contact with) anarray of parallel channels with consistent cross sections and length,each having periodic valve openings such that each can form anindependent series of channels. Valves can be individual or can beformed by collective structures that interact. For example, a perforatedcylinder can have an array of openings leading to channels. A slightlysmaller, concentric non-perforated cylindrical sleeve can be configuredto nest snuggly against the perforated cylinder and slide along theirshared longitudinal axis, thereby allowing fluid to flow through aseries of successively revealed openings, for example. Othercomplimentary structures are also possible, including one cylinder withholes in part of a surface, and a complimentary partial closure cylinderconfigured to cover all, some, or none of the holes, depending on atwist angle of the closure cylinder. By staggering openings to parallelchannels, sliding or twisting such a sleeve or closure structure canprogressively and incrementally expose more or less total channel lengthto cooling fluid flow. Such collective valve structures can providemechanically elegant and finely tunable cooling capabilities. Valves canbe controlled (e.g., actively) or be able to automatically respond tointernal thruster conditions due to their structure. A control systemcan include one or more sensors, one or more processors, and/or one ormore actuators (e.g., a drive, motor, gearbox, mechanical links, etc.).Such a control system can be used to evaluate physical conditions and/ordecide (or receive decisions for) system changes. For example, a controlsystem can cause valves to open or close to allow use of one or morepropellant types and/or states, to control speed or location of arocket, and/or to account for internal or external conditions or goals.A thruster system can thus be internally or remotely controlled. Somevalves can be designed to actively respond to pressures and temperatureswithout external controls. The same or a similar control system cancontrol other moving parts of a rocket. For example, solar reflectors,collectors, modulators, diffusers, concentrators, etc. can have theirangles and surface areas adjusted for better performance. Controlsystems can share or have redundant sensors, and they can respondcollectively. For example, excessive thruster wall heat can be addressedby one or more changes, including valve changes (causing changes tofluid flow), aperture changes (causing changes to solar flux entering achamber), and/or reflector angle changes (causing changes to solarenergy initially collected or relayed).

Embodiments of the described systems can help adjustably cool peripheralstructures for any type of rocket engine, including chemical,electrical, nuclear, and solar thermal rocket engines. For example, arocket engine can include a combustion chamber having a strongperipheral wall, reinforced, thickened, engineered, or otherwiseconfigured to contain propellant and allow for combustion. Aconverging/diverging rocket nozzle at one end of the combustion chambercan be configured to expel propellant to produce rocket thrust. Energycan be introduced into the combustion chamber through electrical,chemical, nuclear, solar, or other mechanisms. At least one coolingchannel can be provided in, adjacent to, or otherwise in thermal contactwith a peripheral wall or other structure of the combustion chamber (orother portion of the rocket, such as an exhaust jacket). One or moreintermediate openings can be provided in a cooling channel or channels.One or more valves can be provided to allow fluid access at theintermediate (and other) openings of the channel or channels. Thesevalves can be configured to change a deployed length of the at least onecooling channel, thereby adjusting a cooling effect (e.g., within astructure of the rocket engine such as the peripheral wall of thethruster, combustion chamber, exhaust, etc.).

FIG. 14 illustrates an embodiment of an off an outward facing solarcollector. In particular, FIG. 14 depicts an off axis concentratingreflector (typically parabolic) 22 and 23 concentrating incomingparallel light 20 into a high intensity beam 3. Light from the sun 20 isfocused to point 24 by a primary reflector 22 and collimated by asecondary reflector 23 into a beam 3. The beam passes through an opening25 in the primary reflector 22 and enters an internal optical assembly26 in the spacecraft 27 to provide power to other systems.

One aspect of the embodiment of FIG. 14 is that the primary reflector 22faces away from the spacecraft 27. As such, a hole 25 is formed in theprimary reflector to allow the concentrated light 3 to reach thespacecraft 27 in a mass and space efficient architecture. In someembodiments, the primary reflector 22 can follow a parabolic curve withthe focus at a point 24. FIG. 14 shows an embodiment in which thereflector 23 and 23 is configured in an off-axis Cassegrainconfiguration, but on-axis configurations are also possible. As the Sunis not infinitely large and infinitely far away, the light from the sunhas a slight divergence (aka finite etendue). To account for thisdivergence, the secondary reflector 23 can be a segment of a hyperboliccurve with one focus coincident with the focus of the primary reflectorat a point 24. The second focus of this hyperbolic curve is locateddownstream in the collimated beam 3 such that it compensates for thelight's divergence.

Benefits of an outward facing primary reflector 22 include safetyagainst pointing errors and a simpler deployment system. As realitylacks perfect parabolas and exact positioning, misalignments areexpected. For example, if the spacecraft 27 temporarily loses theability to point the primary reflector 22 successfully at the Sun andinstead points the primary reflector 22 at an angle somewhat away fromthe Sun, the highly concentrated light 3 from a large reflector could beaccidentally directed to hit parts of the spacecraft that would bedamaged by the highly concentrated light 3. If the high intensity beamof light 3 were to accidentally hit the spacecraft 27, irreparabledamage could occur. By passing the beam 3 back through the hole 25 inthe primary reflector 22, damage caused by misalignments are mitigatedas stray light is harmlessly reflected off the primary reflector 22 backinto space.

Spacecraft 27 often have limited allotted volume in the payload sectionof launch vehicles. The geometry of an off-axis reflector 22 can bestored during launch more simply and efficiently in the given volumethan other embodiments. A simple and efficient storage configuration isreliable and may have less mass, thereby reducing costs. In addition, inan off-axis configuration such as that shown in FIG. 14, practitionerswith reasonable skill in the field will understand that the primaryreflector 22 is often the largest-mass component of the optical systemand it is not positioned far from the spacecraft 27 as in other systems.The close proximity of the primary reflector 22 has a smaller structuralmoment arm and therefor allows a less massive structural support system.As saving mass is beneficial in spacecraft 27 design, this is desirable.

It is of note that solar concentrators can be used to power solarthermal rocket engines but may also have utility on other applicationssuch as solar thermal power systems and so applications of the presentdisclosure are not limited to propulsion but also include thermal andpower systems of other types.

The diagram of FIG. 14 shows just one primary 22 and one secondaryreflector 23. In application a given spacecraft 27 may have multipleprimary reflectors 22, each with their own secondary reflector 23 andeach primary reflector 22 may have multiple secondary reflectors 23. Itis desirable that the primary reflector 22 is directed away from thespacecraft 27 and is located near the spacecraft 27 to afford theaforementioned benefits.

Practitioners with reasonable skill in the field will understand thateach of the optical elements in the system describe herein could becomprised of shapes made from ideal mathematical surfaces such asparabolas, ellipses, and/or hyperbolas, or could have shapes determinedby numerical optimization using modern ray tracing tools, the effect ofwhich is it optimize the design and improve performance, but may notfundamentally alter the salient features of the design.

Example Aspects of Solar Thermal Rockets

In one aspect, a system of decreasing stress from a thermal gradient ina solar thermal rocket, the system comprises: one or more opticalelements configured to collect and concentrate radiant solar energy,focusing the solar energy toward a propulsion chamber; the propulsionchamber configured to contain pressurized propellant gas at a hightemperature within a propulsion chamber wall having a thickness; awindow configured to admit the concentrated solar energy into thepropulsion chamber; a refractory ceramic within the propulsion chamberthat is configured to absorb the concentrated solar energy that entersthrough the window, heat to high temperature, and transfer heat energyto propellant gas flowing adjacent to one or more surfaces of therefractory ceramic; a converging-diverging rocket nozzle configured toexpel high speed gas from the propulsion chamber after it has flowedgenerally through or next to the refractory ceramic, thereby creatingrocket thrust; a jacket formed to generally concentrically surround thepropulsion chamber and rocket nozzle while allowing propellant to flowunimpeded from the nozzle, the jacket and the window having an outersurface that, in use, has a low temperature that is generally the sameas the surrounding space or ambient air; and a heat shield comprising anintermediate concentric layer positioned generally inside the jacket andoutside the propulsion chamber and rocket nozzle, the shield configuredto have an intermediate temperature between the high temperature and thelow temperature, thereby reducing stress on the propulsion chamberresulting from a thermal gradient across the thickness of the propulsionchamber wall.

The refractory ceramic within the propulsion chamber comprise anopen-cell ceramic foam configured to permit a propellant gas to flowthrough the ceramic foam and absorb heat by conduction due to contactwith the large internal surface area of the ceramic foam.

The propulsion chamber and converging-diverging rocket nozzle can befabricated as a continuous hollow cylindrical structure and thepropulsion chamber can be fabricated from a refractory ceramic capableof high operating temperatures. In some embodiments, the continuoushollow cylindrical structure and the propulsion chamber can be formed ofother materials, such as one or more metals.

The heat shield can be spaced from the propulsion chamber and from themetal jacket by an empty space that is open to ambient such that duringspace flight it will be evacuated.

Another aspect is a solar thermal rocket engine configured to use singlyor in combination simultaneously one or more liquid or gaseouspropellant fluids.

The propellant flow rates through the vaporizer (regenerativeheating/cooling channels) can be dynamically adjusted to accommodatechanges in propellant types while maintaining engine operatingtemperatures within design limits.

The vaporizer (regenerative heating/cooling channels) can be providedwith dynamically adjustable length and cross section area to control andadjust the surface area available for heat transfer to and frompropulsion fluids. In some cases, sections of the channels can bebypassed in for use with propellants that are more easily vaporized. Thebypass can be accomplished with valves that direct the propellantthrough tube lengths that circumvent more circuitous channels.

The vaporizer channels of fixed length can be provided with multipleinlet ports along the length of the channel and multiple inlet valvesfor injecting propellant fluids into adjustable lengths of channel.

The multiple parallel vaporizer channels of fixed length can be eachprovided with inlet valves to independently adjust the flow ofpropulsion fluids through a multiplicity of channels.

The valves can be used to admit various propellant fluids singly or incombination into heat exchangers and to adjust their flows toaccommodate changing thermodynamic requirements and to control pressureinstabilities during different phases of rocket operations includingstartup, shutdown, and continuous operation at variable thrust levels.

The adjustable regenerative channels can have a helical form (e.g.,wound around the surface of or manufactured into the pressure chamberand exhaust cone of a rocket engine) or of axial form (e.g., bondedalong the length of or manufactured into the rocket engine). Multipleform can be combined to allow for greater flexibility and variation, forexample, depending on which channels are opened for use.

The collected solar energy can be concentrated and focused through atransparent pressure-resistant window into a solar absorbing structure.This can comprise multiple surfaces that interact with energy. Forexample, it can comprise a labyrinthine assembly of multiple partiallyreflecting, partially transmitting, and partially absorbing surfaceswhere solar radiant energy is converted to thermal heat energy in theabsorbing surfaces. The assembly can further contain pressurizedpropulsion gas in the spaces between absorbing surfaces, such thatthermal heat is transferred to the propulsion gas by both thermalradiation and by direct gaseous fluid conduction from the absorbingsurfaces.

The pressurized and heated propulsion gas can be expelled through aconverging/diverging rocket nozzle to produce rocket thrust.

The assembly of surfaces can be constructed from metallic alloys and/orceramics, for example.

The engine can have an assembly of surfaces that comprises reflectingand absorbing passages of triangular, rectangular, hexagonal, orcircular cross-sections.

The engine can further comprise a solar flux modulating device. Forexample, this device can be an optical aperture of adjustable openingarea that is disposed between a solar collector and a solar absorber forthe purpose of dynamically adjusting the total thermal power input tothe rocket engine. A coating, shutter or shield feature can also oralternatively be used to adjust solar flux or alter a balance betweenreflection, absorption, and transmission (see FIG. 6C, for example).

The engine can maintain a store of a cleaning propellant and can beconfigured to use the cleaning propellant periodically during operationto clean the inside surface of the engine and remove deposits madethrough the operation of the engine when using propellants that leavesuch deposits behind including carbon depots left behind by coking, theengine configured to operate with different combinations of depositingpropellants and cleaning propellants.

Yet another aspect is a rocket engine system comprising: a firstpropellant container configured to hold deposit forming propellant; asecond propellant container configured to hold deposit cleaningpropellant; a passage connected to both the first and second propellantcontainers; a manifold in fluid communication with the passage andconfigured to select or combine propellant from the containers; and acleaning control system configured to control the manifold, therebyusing the two propellants in combination to reduce deposits within therocket engine system.

The rocket engine can further comprise: a third container configured tohold cooling fluid; a passage configured to direct cooling fluid throughstructural elements of the rocket engine; and a cooling control systemconfigured to control the flow of the cooling fluid, thereby at leastperiodically cooling to maintain structural integrity of the rocketengine. Conclusion

Unless the context clearly requires otherwise, throughout thedescription and the claims, the words “comprise,” “comprising,” and thelike are to be construed in an inclusive sense, as opposed to anexclusive or exhaustive sense; that is to say, in the sense of“including, but not limited to.” The word “coupled”, as generally usedherein, refers to two or more elements that may be either directlyconnected, or connected by way of one or more intermediate elements.Likewise, the word “connected”, as generally used herein, refers to twoor more elements that may be either directly connected, or connected byway of one or more intermediate elements. Additionally, the words“herein,” “above,” “below,” and words of similar import, when used inthis application, shall refer to this application as a whole and not toany particular portions of this application. Where the context permits,words in the above Detailed Description using the singular or pluralnumber may also include the plural or singular number respectively. Theword “or” in reference to a list of two or more items, that word coversall of the following interpretations of the word: any of the items inthe list, all of the items in the list, and any combination of the itemsin the list.

Moreover, conditional language used herein, such as, among others,“can,” “could,” “might,” “can,” “e.g.,” “for example,” “such as” and thelike, unless specifically stated otherwise, or otherwise understoodwithin the context as used, is generally intended to convey that certainembodiments include, while other embodiments do not include, certainfeatures, elements and/or states. Thus, such conditional language is notgenerally intended to imply that features, elements and/or states are inany way required for one or more embodiments or that one or moreembodiments necessarily include logic for deciding, with or withoutauthor input or prompting, whether these features, elements and/orstates are included or are to be performed in any particular embodiment.

The above detailed description is not intended to be exhaustive or tolimit the invention to the precise form disclosed above. While specificembodiments of, and examples are described above for illustrativepurposes, various equivalent modifications are possible within the scopeof the disclosed invention(s), as those skilled in the relevant art willrecognize. For example, while processes or blocks are presented in agiven order, alternative embodiments may perform routines having steps,or employ systems having blocks, in a different order, and someprocesses or blocks may be deleted, moved, added, subdivided, combined,and/or modified. Each of these processes or blocks may be implemented ina variety of different ways. Also, while processes or blocks are attimes shown as being performed in series, these processes or blocks mayinstead be performed in parallel, or may be performed at differenttimes.

The teachings provided herein can be applied to other systems, notnecessarily the system described above. The elements and acts of thevarious embodiments described above can be combined to provide furtherembodiments.

While certain embodiments have been described, these embodiments havebeen presented by way of example only, and are not intended to limit thescope of the disclosure. Indeed, the novel methods and systems describedherein may be embodied in a variety of other forms; furthermore, variousomissions, substitutions and changes in the form of the methods andsystems described herein may be made without departing from the spiritof the disclosure. The accompanying claims and their equivalents areintended to cover such forms or modifications as would fall within thescope and spirit of the disclosure.

What is claimed is:
 1. A solar thermal rocket engine, comprising: asolar thermal thruster configured to: receive solar energy and one ormore propellants, and heat the one or more propellants using the solarenergy to generate thrust, wherein the solar thermal thruster is furtherconfigured to adapt for a plurality of different propellant types,either singly or in combination simultaneously, and wherein the solarthermal thruster is further configured to use one or more of thepropellant types in both liquid and gaseous states.
 2. The engine ofclaim 1, wherein the solar thermal thruster comprises: one or moreregenerative channels configured to direct flow of and simultaneouslychange thermal energy of the one or more propellants.
 3. The engine ofclaim 2, further configured to adjust a length and cross sectional areaof the regenerative channels.
 4. The engine of claim 2, wherein: each ofthe regenerative channels has a fixed total length and comprises aplurality of inlet ports arranged along that length, and the enginefurther comprises a plurality of inlet valves associated with the inletports and configured to selectively inject the one or more propellantsinto the regenerative channels via selected inlet ports such that adeployed length of the regenerative channels is adjustable.
 5. Theengine of claim 2, wherein: each of the regenerative channels has afixed total length, includes a plurality of inlet ports, and the enginefurther comprises a plurality of inlet valves configured to block,permit, or change flow of the one or more propellants into the inletports, thereby adjusting a flow of the one or more propellants throughthe regenerative channels.
 6. The engine of claim 2, wherein: the solarthermal thruster comprises a pressure chamber and an exhaust cone, thepressure chamber connected to and axially aligned with the exhaust cone;and the regenerative channels follow a helical path around or through asurface of at least one of the pressure chamber and the exhaust cone. 7.The engine of claim 1, further comprising: one or more electric thermalelements configured to adjust a temperature of the solar thermalthruster; one or more heat exchangers; and one or more valves to admitthe one or more propellants into the heat exchangers and adjust flows ofthe one or more propellants to accommodate changing thermodynamiccharacteristics during different phases of rocket operations includingstartup, shutdown, and continuous operation at variable thrust levels.8. The engine of claim 1, wherein the solar thermal thruster comprises:a solar absorbing structure comprising an assembly of partiallyreflecting, partially transmitting, and partially absorbing surfaces,thereby converting solar energy into thermal energy within the surfaces,and a transparent pressure-resistant window configured to transmit thesolar energy into the solar absorbing structure and contain the gaseswithin the chamber.
 9. The engine of claim 8, wherein the solarabsorbing structure is further configured to contain the one or morepropellants in spaces between the surfaces such that the thermal heatenergy is transferred to the one or more propellants via both thermalradiation and fluid conduction from the surfaces.
 10. The engine ofclaim 8, wherein the solar absorbing structure of surfaces is formed ofone or more of the following materials: metallic alloys and ceramics.11. The engine of claim 8, wherein the solar absorbing structurecomprises a plurality of reflecting and absorbing passages havingtriangular, rectangular, hexagonal, octagonal, or circularcross-sections, the reflecting and absorbing passages further configuredto provide a volume over which the solar energy is absorbedsimultaneously with providing a surface area and shape of a surface overwhich the propellant flows to absorb heat from the solar absorberstructure.
 12. The engine of claim 11, wherein the solar absorberfurther comprises a multi-surface light trapping solar absorbercomprising a honeycomb lattice, a bundle of thin walled tubes, or a coilof thin sheets.
 13. The engine of claim 1, wherein the solar thermalthruster further comprises one or more peripheral cooling channels, andthe thruster adapts for different propellant types by using at least onevalve to adjust a deployed length of at least one cooling channel,thereby adjusting a cooling effect in the thruster.
 14. The engine ofclaim 1, further comprising: a solar collector configured to collect andredirect the solar energy; a solar absorber configured to absorb atleast a portion of the solar energy; and an adjustable solar fluxmodulator disposed between the solar collector and the solar absorber,the modulator configured to adjust thermal power input to the solarthermal thruster.
 15. The engine of claim 14, wherein the solar fluxmodulator comprises one or more of a variable geometry aperture, ashuttered opening, one or more blinds, and an opening at least partiallycovered with a material having variable optical transmissivity.
 16. Theengine of claim 1, further comprising: a store of cleaning propellant,wherein the solar thermal thruster is further configured to periodicallyuse the cleaning propellant to clean an inside surface of the solarthermal thruster and remove deposits made through the operation of thesolar thermal thruster when using the one or more propellants.
 17. Theengine of claim 15, wherein the solar thermal thruster is furtherconfigured to operate with different combinations of depositingpropellants and cleaning propellants.
 18. The engine of claim 1, whereinthe window in the thruster is configured as a lens to furtherconcentrate the sunlight into the thruster so as to increase the peaktemperature inside the thruster and increase the performance of thethruster.
 19. An adjustable jacket cooling system for a rocket engine,the system comprising: a combustion chamber having a strong peripheralwall and configured to contain propellant; a converging/diverging rocketnozzle at one end of the combustion chamber and configured to expelpropellant to produce rocket thrust; at least one cooling channel inthermal contact with the peripheral wall; at least one intermediateopening in the cooling channel; at least one valve configured to changea deployed length of the at least one cooling channel, thereby adjustinga cooling effect for the peripheral wall.
 20. The system of claim 19,wherein the cooling channel and valve are configured to use theintermediate opening to change the structural cooling capability of therocket engine such that it can operate effectively using at least twodifferent types of propellant that have significantly differentthermodynamic properties.
 21. The adjustable jacket cooling system ofclaim 19 adapted for use with a solar thermal rocket engine comprising:a solar thermal thruster configured to receive solar energy and one ormore propellants, and heat the one or more propellants using the solarenergy to generate thrust, wherein the adjustable jacket cooling systemallows the thruster to use a plurality of different propellant types inone or both of their liquid and gaseous states.
 22. The cooling systemand engine of claim 21, wherein the at least one cooling channelcomprises one or more regenerative channels configured to direct flow ofand simultaneously change thermal energy of the one or more propellants.23. The cooling system and engine of claim 22, wherein the adjustablejacket cooling system is further configured to adjust a deployed lengthof the one or more regenerative channels.
 24. A solar concentratorconfigured to power the solar thermal rocket engine of claim 1, thesolar concentrator comprising: a primary reflector; and a secondaryreflector, wherein the primary reflector is configured to concentratethe solar energy towards the secondary reflector, and wherein thesecondary reflector is configured to reflect the solar energy into aless converging, slightly diverging, or parallel beam, the intensity ofthe reflected solar energy being greater than the solar energy prior tobeing reflected by the secondary reflector.
 25. The solar concentratorof claim 24, wherein two reflectors are arranged in a Cassegrainconfiguration.
 26. The solar concentrator of claim 24, wherein theprimary reflector comprises an orifice, and wherein the primaryreflector and the secondary reflector are oriented such that thereflected solar energy passes through the orifice in the primaryreflector during normal operation prior to entering the solar thermalrocket engine.
 27. The solar concentrator of claim 26, wherein theorifice is positioned and configured to mitigate damaging effects ofpointing errors by rejecting the reflected solar energy when incorrectlypointed.
 28. A rocket engine system comprising: a first propellantcontainer configured to hold deposit-forming propellant; a secondpropellant container configured to hold deposit-cleaning propellant; apassage connected to both the first and second propellant containers; amanifold in fluid communication with the passage and configured toselect or combine propellant from the containers; and a cleaning controlsystem configured to control the manifold, thereby using propellants toreduce deposits within the rocket engine system.
 29. The rocket enginesystem of claim 28, further comprising: a third container configured tohold cooling fluid; a passage configured to direct cooling fluid throughstructural elements of the rocket engine; and a cooling control systemconfigured to control the flow of the cooling fluid, thereby at leastperiodically cooling to maintain structural integrity of the rocketengine.